The invention relates to a method and a device for aligning a space vehicle, particularly a geostationary satellite, in a reference direction.
Space vehicles, such as communication satellites, must take up a defined alignment with respect to the earth and/or the sun in certain operating stages. For this purpose, the satellite must be aligned with respect to the sun by means of a specific direction set in the satellite system of coordinates, the so-called reference direction. For this reason, two generally orthogonal measurements have been made up to now in the case of geostationary satellites, by means of which the momentary direction of the sun is determined. The deviation of this momentary direction of the sun from the reference direction is then used for aligning the satellite by means of a control system. In the control system, control signals are computed from the measured data and other characteristics of the satellite and of the sensors and are converted into control signals which are then supplied to the actuator elements of the satellite.
For the previous method for aligning the satellite with respect to the sun, at least two sensors have always been required, an additional sensor being present as a result of redundancy requirements. In addition, the mentioned earth acquisition can be carried out by means of the conventional method only when certain marginal geometric conditions exist, specifically when the sun, the satellite and the earth take up a defined constellation with respect to one another. If, for example, the earth, as a result of disturbances disappears from the catch area of the satellite and its sensors, so that the so-called earth reference is lost for the satellite, considerable waiting periods must sometimes be accepted until the mentioned defined constellation is reached and an earth acquisition becomes possible.
It is an object of the invention to provide a method and a device of the above-mentioned type which permit an alignment of the satellite in the direction of a reference object, particularly the sun, by means of lower technical expenditures and a smaller number of measurements, although a simple control law is made available which is not inferior to the known control law.
According to the invention this object is achieved by providing a device and method for aligning a space vehicle with respect to a reference object, including obtaining a measuring direction by determining the angular orientation direction of the space vehicle with respect to the reference object and subsequently aligning the space vehicle by means of actuators in such a manner that the measuring direction corresponds to a reference direction on the space vehicle, said actuators being acted upon by control signals which are derived from the measuring direction to control torques on the space vehicle, wherein the orientation of space vehicle and reference object is determined only with respect to a single measuring direction along a main axis of a direction sensor, and wherein the following steps are taken if the measuring direction does not correspond to the reference direction on the space vehicle:
a) the space vehicle is rotated around the reference direction on the space vehicle,
b) a fault component of the direction of the reference object situated in the measuring direction is controlled by locking on a first control torque perpendicularly with respect to the measuring direction and perpendicularly with respect to the reference direction on the space vehicle,
c) the unmeasurable component of the direction of the reference object situated perpendicularly with respect to the measuring direction is controlled by locking on a second control torque perpendicularly with respect to the reference direction and perpendicularly with respect to the first steering moment.
According to certain preferred embodiments, the first control torque is represented as a vector in such a manner that EQU a.sub.2 =S.sub.R e.sub.M (N.sub.sy -N.sub.by) or (5) EQU a.sub.2 =Oe.sub.M (N.sub.sy -N.sub.by)
wherein e.sub.M is the vector of the measuring direction S.sub.R is the cross product of the vector of the reference direction with the vector of the measuring direction, N.sub.sy is the component of the measured direction of the sun (S.sub.b) in the measuring direction, N.sub.by is a preset basic value of this component, and O is the optical axis of a sensor determining the direction of the sun.
According to certain preferred embodiments, the second control torque is represented as a vector in such a manner that EQU a.sub.3 =S.sub.R (S.sub.R e.sub.m)(N.sub.sy -N.sub.by) or (7) EQU a.sub.3 =S.sub.R Oe.sub.M (N.sub.sy -N.sub.by) or a.sub.3 =(-e.sub.M)(N.sub.sy -N.sub.by)
wherein e.sub.M is the vector of the measuring direction, S.sub.R is the cross product of the vector of the reference direction with the vector of the measuring direction, N.sub.sy is the component of the measured direction of the sun (S.sub.b) in the measuring direction, N.sub.by is a preset basic value of this component, and O is the optical axis of a sensor determining the direction of the sun.
According to certain preferred embodiments, the following control law is used: EQU u=-K.sub.D (w-cS.sub.R)+K.sub.p (S.sub.R +S.sub.R S.sub.R.sup.T -I)e.sub.M L(N.sub.sy -N.sub.by) (8)
wherein u are the input signals for the actuators, K.sub.D are the control parameters for the rotational speeds around the space vehicle axes, K.sub.p are the control parameters for the position of the space vehicle, w are the rotational speeds of the space vehicle around are space vehicle axes, c is the value for the rotating rate of the space vehicle, S.sub.R is the vector of the reference direction, S.sub.R is the cross-product of the reference vector S.sub.R with the vector of the measuring direction e.sub.M, S.sub.R.sup.T is the transposed vector of the reference direction, I is the unity matrix, L is a limiter N.sub.sy is the measured component of the direction S.sub.b of the sun in the measuring direction, and N.sub.by is a preset basic value for this component.
In a preferred embodiment of the apparatus according to the present invention, an arrangement is provided:
wherein a sun sensor is provided, the output signal of which, after a zero point correction, is supplied to a limiter,
wherein, in a first branch, a first multiplier is provided for calculating the rotating direction of a control vector (a.sub.2) from the measuring direction (e.sub.M) and a reference direction (O or SR),
wherein, in a second branch, another multiplier is provided for calculating the rotating direction of another control vector also from the measuring direction (e.sub.M) and a reference direction (S.sub.R or O),
wherein the rotating direction vectors are each multiplied with the output signal of the limiter and are subsequently added to one another,
wherein the sum signal is supplied to a multiplier for multiplication with a control parameter (K.sub.p),
wherein, in a third branch, a multiplier is provided for calculating a rotational speed vector (w.sub.b) of the space vehicle around three axes on the space vehicle from a reference direction (S.sub.R) in a constant (c),
wherein three multipliers are provided in which the sum signal combined from the two control vectors (a.sub.2, a.sub.3), the rotational speed vector (w.sub.b) and the rotational speeds (w) of the space vehicle are multiplied with control parameters (K.sub.p, K.sub.D),
and wherein the output quantities of these multipliers are supplied to an adding device, the output signal of which is the control signal.
In the case of the suggested method of the invention, only a single measuring signal is required with respect to a direction; i.e., also only a single sensor by means of which the direction of the reference object, in the following generally called the sun, is determined with respect to the measuring direction of the sensor, i.e., its main axis. According to this control concept, when there is a deviation from the alignment with the sun, the space vehicle is first rotated around an axis which coincides with the reference direction on the space vehicle. The actuators are driven by control signals in such a manner that the component of the momentary direction of the sun is controlled in the measuring direction. The second component of the momentary direction of the sun which is perpendicular with respect to the measuring direction and which cannot be measured, subsequently is also controlled in that control signals are supplied to the actuators in such a manner that the control torques act upon the space vehicle perpendicularly with respect to the reference direction and perpendicularly with respect to the first control torque. For this maneuver, the momentary direction of the sun is made to conform with the reference direction.
The two control torques may either be geared to the reference direction on the space vehicle or, particularly if the sensors, such as sun sensors, have fields of view with a directional angle of more than 90.degree., may use the optical axis of this sensor as a reference. As a result, convergency difficulties are avoided for certain constellations for the direction of the sun and the reference direction. This has the additional advantage that the vectors occurring in the control law are constant for the individual sensors; i.e., do not depend on the reference direction which means that the calculation of the control law is simplified.
By means of a method and a device according to the invention, at least two sun sensors may be saved, for example, in the case of communication satellites, specifically the generally primary sensor and the redundant sensor, by means of which, according to the conventional method, the second direction components of the sun vector pointing to the sun are measured corresponding to the direction of the sun and which are required only for maneuvers for the earth acquisition.
In addition, according to the invention, the waiting period for the earth acquisition can be reduced after the loss of the earth reference. Since only one sensor is required, such an earth acquisition, after the loss of the earth reference, is also possible by means of additional uniaxial sun sensors existing for other maneuvers. Thus, for a new earth acquisition, there has to be no waiting for the constellation between the sun, the space vehicle and the earth which is defined in the known methods.
Although the method according to the invention can be carried out with considerably reduced technical measuring expenditures, the time periods which are required for an alignment of the space vehicle are no longer than previously.
Other objects, advantages and novel features of the present invention will become apparent from the following detailed description of the invention when considered in conjunction with the accompanying drawings.